Many rockets and missiles include one or more solid rocket motors to generate thrust to achieve and/or maintain flight, and/or to control in-flight direction. A solid rocket motor may include, for example, a motor case and a nozzle. Typically, the motor case defines a combustion chamber, in which propellant is loaded and combusted to generate high-energy combustion gas. The nozzle is typically in fluid communication with the combustion chamber and thus receives the high-energy combustion gas. The nozzle may include a convergent inlet section, a divergent outlet section, and an interposing nozzle throat. Combustion gas generated in the combustion chamber flows through the nozzle, generating a thrust.
Solid rocket motors, such as the one briefly described above, are used in both strategic and tactical rockets and missiles. In general, strategic missiles are used for long duration missions, whereas tactical missiles are used for relatively short duration flight missions. Thus, many tactical missiles use solid rocket motors with relatively high burn rates. If the burn rate of the solid rocket motor can be controlled, the thrust generated by the motor can also be controlled, and the overall operation of the rocket can become more efficient.
As is generally known, the burn rate of the propellant in a solid rocket motor may vary with the pressure in the combustion chamber. For example, if the combustion chamber pressure increases, the propellant burn rate increases, and the thrust generated by the rocket motor will concomitantly increase. Conversely, if the combustion chamber pressure decreases, the propellant burn rate decreases, and the thrust generated by the rocket motor decreases. One way of controlling combustion chamber pressure, and thus propellant burn rate, is by controlling the effective flow area of the nozzle throat. For example, if the effective flow area of the nozzle throat decreases, combustion chamber pressure increases, and vice-versa.
Various systems and methods have been developed for varying the effective flow area of a rocket nozzle throat. Such systems and methods include selectively venting combustion gas from the combustion chamber, controlling overall exhaust flow, and selectively physically altering the flow area nozzle throat. Although these systems and methods are effective, each suffers certain drawbacks. For example, the present systems and methods can significantly affect overall rocket motor efficiency, and may rely on fairly complex and costly components and control systems.
Hence, there is a need for a system and method of controlling rocket motor nozzle effective flow area that does not significantly affect overall motor efficiency and/or does not rely on fairly complex and costly components and control systems. The present invention addresses one or more of these needs.